scholarly journals Numerical simulation of the interaction of the disturbed boundary layer with an incident shock

2021 ◽  
Vol 2057 (1) ◽  
pp. 012005
Author(s):  
D V Khotyanovsky ◽  
A N Kudryavtsev ◽  
A I Kutepova
Keyword(s):  
Boundary Layer ◽  
Separation Zone ◽  
Boundary Layer Separation ◽  
Incident Shock ◽  
Supersonic Boundary Layer ◽  
Linear Stability Theory ◽  
Oblique Shock Wave ◽  
Separation Flow ◽  
Turbulent Transition ◽  
Unstable Boundary Layer

Abstract Interaction of the disturbed supersonic boundary layer with an incident oblique shock wave is studied numerically with eddy-resolving numerical simulations. Eigenmodes of the linear stability theory are used to generate the inflow boundary layer disturbances. The evolution of unstable boundary-layer disturbances, effects of the incident shock on the disturbances, effects of the disturbances on the boundary layer separation, flow dynamics in the separation zone, and laminar-turbulent transition are studied.

scholarly journals The influence of moderate angle-of-attack variation on disturbances evolution and transition to turbulence in supersonic boundary layer on swept wing

2019 ◽  
Vol 234 (1) ◽  
pp. 96-101
Author(s):  
Alexander Kosinov ◽  
Nikolai Semionov ◽  
Yury Yermolaev ◽  
Boris Smorodsky ◽  
Gleb Kolosov ◽  
...  
Keyword(s):  
Boundary Layer ◽  
Theoretical Study ◽  
Angle Of Attack ◽  
Reynolds Numbers ◽  
Supersonic Boundary Layer ◽  
Linear Stability Theory ◽  
Transition To Turbulence ◽  
Computational Results ◽  
Swept Wing ◽  
Turbulent Transition

The paper is devoted to an experimental and theoretical study of effect of moderate angle-of-attack variation on disturbances evolution and laminar-turbulent transition in a supersonic boundary layer on swept wing at Mach 2. Monotonous growth of the transition Reynolds numbers with angle of attack increasing from −2° to 2.7° is confirmed. For the same conditions, calculations based on linear stability theory are performed. The experimental and computational results show a favourable comparison.

Combined influence of coating permeability and roughness on supersonic boundary layer stability and transition

2016 ◽  
Vol 798 ◽  
pp. 751-773 ◽  
Author(s):  
V. I. Lysenko ◽  
S. A. Gaponov ◽  
B. V. Smorodsky ◽  
Yu. G. Yermolaev ◽  
A. D. Kosinov ◽  
...  
Keyword(s):  
Boundary Layer ◽  
Plate Boundary ◽  
Leading Edge ◽  
Quantitative Agreement ◽  
Supersonic Boundary Layer ◽  
Linear Stability Theory ◽  
Porous Coating ◽  
Boundary Layer Stability ◽  
Turbulent Transition ◽  
Laminar Turbulent Transition

A joint theoretical and experimental investigation of the influence of the surface permeability and roughness on the stability and laminar–turbulent transition of a supersonic flat-plate boundary layer at a free-stream Mach number of $M_{\infty }=2$ has been performed. Good quantitative agreement of the experimental data obtained with artificially generated disturbances performed on models with various porous inserts and calculations based on linear stability theory has been achieved. An increase of the pore size and porous-coating thickness leads to a boundary layer destabilization that accelerates the laminar–turbulent transition. It is shown that as a certain (critical) roughness value is reached, with an increase in the thickness of the rough and porous coating, the boundary layer stability diminishes and the laminar–turbulent transition is displaced towards the leading edge of the model.

An Automated CFD Design and Analysis Tool for Inlet-Bleed Systems

1998 ◽  
Author(s):  
Tom I-P. Shih ◽  
Yu-Liang Lin ◽  
Andrew J. Flores ◽  
Mark A. Stephens ◽  
Mark J. Rimlinger ◽  
...  
Keyword(s):  
Boundary Layer ◽  
Three Dimensional ◽  
Incident Shock ◽  
Navier Stokes ◽  
Oblique Shock Wave ◽  
Design Parameters ◽  
Analysis Tool ◽  
Cfd Analysis ◽  
Circular Holes ◽  
Embedded Grids

Abstract A pre-processor was developed to assist CFD experts and non-experts in performing steady, three-dimensional Navier-Stokes analysis of a class of inlet-bleed problems involving oblique shock-wave/ boundary-layer interactions on a flat plate with bleed into a plenum through rows of circular holes. With this pre-processor, once geometry (e.g., hole dimensions and arrangement) and flow conditions (e.g., Mach number, boundary-layer thickness, incident shock location) are inputted, it will automatically generate every file needed to perform a CFD analysis from the grid system to initial and boundary conditions. This is accomplished by accessing a knowledge base established by experts who understand both CFD and the class of problems being analyzed. For experts in CFD, this tool greatly reduces the amount of time and effort needed to setup a problem for CFD analysis. It also provides experts with knobs to make changes to the setup if desired. For non-experts in CFD, this tool enables reliable and correct usage of CFD. A typical session on a workstation from data input to the generation of all files needed to perform a CFD analysis involves less than ten minutes. This pre-processor, referred to as AUTOMAT-V2, is an improved version of a code called AUTOMAT. Improvements made include: (1) multi-block structured grids can be patched in addition to being overlapped; (2) embedded grids can be introduced near bleed holes to reduce the number of grid points/cells needed by a factor of up to four; (3) grid systems generated allow up to three levels of multigrid; (4) CFL3D is supported in addition to OVERFLOW, two well-known and highly regarded Navier-Stokes solvers developed at NASA’s Langley and Ames Research Centers; (5) all files needed to run RONNIE for patched grids and MAGGIE for overlapped grids are also generated; and (6) more design parameters can be investigated including the study of micro bleed and effects of flow/hole misalignments.

Control of incident shock-induced boundary-layer separation using steady micro-jet actuators at M∞ = 3.5

2018 ◽  
Vol 233 (4) ◽  
pp. 1284-1306 ◽  
Author(s):  
Manisankar Chidambaranathan ◽  
Shashi B Verma ◽  
Ethirajan Rathakrishnan
Keyword(s):  
Boundary Layer ◽  
Pitch Angle ◽  
Vortex Generator ◽  
Boundary Layer Separation ◽  
Incident Shock ◽  
Skew Angle ◽  
Zero Crossing ◽  
Shock Generator ◽  
Jet Actuators ◽  
Injection Pressures

Experiments were carried out to control an incident shock-induced separation associated with 22° shock generator in a Mach 3.5 flow using an array of steady micro-jet actuators. Four micro-jet actuator configurations based on the variation in their pitch angle [Formula: see text], skew angle [Formula: see text] and span-wise spacing were used. Each of these configurations were placed 14 δ upstream of the interaction and operated with injection pressures ( Poj) varying from 140 to 643 kPa. While no major variations in separation characteristics were observed for Poj < 140 kPa, significant modifications were observed beyond [Formula: see text] of 140 kPa and until 208.5 kPa. Amongst all the four control configurations, micro-jet vortex generator 2 ([Formula: see text] showed the best control with a 2 δ downstream shift in separation point location relative to no-control. The shift is also accompanied with a change in maximum zero-crossing frequency towards higher frequency (almost twice), a reduction in the intermittency length and an increase in the correlation value between the boundary layer just upstream of the interaction and the intermittent region. These results indicate that the effectiveness of micro-jet vortex generator 2 is probably due to the improved entrainment levels in the shear layer induced by the micro-vortices which are generated downstream of these devices. The increase of the skew angle [Formula: see text] from 180° to 270° for the same pitch angle of β =  45° (micro-jet vortex generator 3) seems to have no major impact on the separation characteristics. The reduction in the span-wise spacing (micro-jet vortex generator 4) resulted in deterioration of the flow field due to the jet-to-jet interaction with increasing injection pressures.

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